The invention relates to gas turbine engines. More particularly, the invention relates to the casting of gas turbine engine blades.
Heat management is an important consideration in the engineering and manufacture of turbine engine blades. Blades are commonly formed with a cooling passageway network. A typical network receives cooling air through the blade platform. The cooling air is passed through convoluted paths through the airfoil, with at least a portion exiting the blade through apertures in the airfoil. These apertures may include holes (e.g., “film holes”) distributed along the pressure and suction side surfaces of the airfoil and holes at junctions of those surfaces at leading and trailing edges. Additional apertures may be located at the blade tip. In common manufacturing techniques, a principal portion of the blade is formed by a casting and machining process. During the casting process a sacrificial core is utilized to form at least main portions of the cooling passageway network.
In turbine engine blades (especially high pressure turbine (HPT) section blades), thermal fatigue of tip region of a blade airfoil is one area of particular concern. U.S. Pat. No. 6,824,359 discloses cooling air outlet passageways fanned along a trailing tip region of the airfoil. U.S. Pat. No. 7,059,834 discloses direction of air through a relief in a wall of a tip pocket to cool a trailing tip portion. U.S. patent application Ser. No. 11/317,394 discloses use of a tip flag passageway to deliver a high volume of cooling air to a trailing tip portion.